FIG. 4 shows a typical three shaft gas turbine engine 10. The gas turbine engine 10 includes an air intake 12, a fan 14 having rotating blades 16, a bypass duct 18 and an engine core 20. The engine core 20 includes an intermediate pressure compressor 22, a high pressure compressor 24, a combustor 26, a turbine arrangement comprising a high pressure turbine 28, an intermediate pressure turbine 30, a low pressure turbine 32 and an exhaust nozzle 34. Air entering the intake 12 is accelerated by the fan 14 and directed into two air flows. The first air flow passes into the engine core 20, and the second air flows along the bypass 18 to provide propulsive thrust.
The engine core air flow travels through the intermediate 22 and high 24 pressure compressors in turn. The compressed air exhausted from the high pressure compressor 24 is mixed with fuel and burnt in the combustor 26. The hot gas expands through and drives the high 28, intermediate 30 and low 32 pressure turbines before being exhausted through the nozzle 34 and adding to the propulsive thrust created by the first air flow. The high 28, intermediate 30 and low 32 pressure turbines respectively drive the high 24 and intermediate 22 pressure compressors and the fan 14 via respective shafts 36, 38, 40.
It is well known that to maintain an efficient gas turbine engine the gap between fan blade tips and the engine casing is closely controlled to minimise the leakage of compressed air over the blade tips and back upstream. To this end, the engine casings often include an attrition or abradable liner which provides a close fitting seal with the blade tips. The abradable liner is initially installed so as to be in contact with the fan blade tips such that the liner is scored by the rotating fan (or compressor as the case may be) during the first few rotations which removes enough material to allow a close fitting free rotation of the blades.
However, during normal engine use the radial position of the rotating blade tips move due to, for example, centrifugal forces, thermal expansion and vibration, and also during harsh operating conditions such as heavy landings or sharp manoeuvres.
This can cause in-service damage to the attrition liner, which, in severe cases, can erode large arcuate sections which then require replacement. Replacement of the liners is expensive both in terms of overhaul cost and the associated loss of service of the engine.
The present invention seeks to provide a solution to help monitor and control attrition liner damage.